AAE 454 Fall 2000 Term Paper

Ultra-High Temperature Ceramics

Written by:Kacie Burton

Introduction

Ultra-High Temperature Ceramics (UHTC’s) have been developed in order to

improve the performance of aerospace vehicles. UHTC’s will allow for new capabilities

of aerospace vehicles due to the fact that they can be used at temperatures up to 5000F.

This is 2000F higher than the hottest areas on the Space Shuttle. Drastically new shapes

in aerospace vehicles will be possible including sharp leading edges as opposed to the

blunt leading edges used on Apollo and currently being used on the Shuttle (Figure 1).

Using UHTC’s the leading edge could be on the order of tenths or hundredths of an inch

rather than feet. Blunt leading edges were used on Apollo and the Shuttle because of the

extreme heat of re-entry, heat that a sharp leading edge made of materials at that time

would not be able to withstand. New re-entry vehicles could now be designed that would

be similar to supersonic aircraft, creating a new category of aerospace vehicles called

sharp bodies.

Ultra-High Temperature Ceramics are made up of a small group of diboride

ceramic matrix composites or CMC that was developed over thirty years ago in U.S. Air

Force sponsored programs. The main CMC’s developed were ZrB2/SiC, HfB2/SiC and

ZrB2/SiC/C because of their unique thermal properties (Clougherty, et.al., 1969).

Arc jet testing has been conducted on UHTC’s at Ames Research Center in order

to determine material performance in an aeroheating environment (Selvaduray, 1997).

Flight-testing on UHTC’s began in May of 1997. Most recently, flight tests were

conducted this past September. The main objectives for these flight tests focused on

performance and life cycle cost. More specifically the flight tests are to demonstrate the

non-ablating performance of the 3-D UHTC sharp nose tip (Figure 2). Testing has also

been conducted in the ground-based arc jet facilities.

Ceramic Matrix Composites

Ultra-High Temperature Ceramics began with the development of diboride

ceramic matrix composites in the late 1960s by the U.S. Air Force. These materials

possess unique thermal properties that are applicable to re-entry bodies at hypersonic

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speeds. The composites that were developed are ZrB2/SiC, HfB2/SiC and ZrB2/SiC/C.

These composites are considered the baseline UHTC’s.

The diborides that the Air Force considered when developing these composites

were zirconium and hafnium because of their thermo-chemical stability and oxidation

resistance at high temperatures (Clougherty, et.al., 1969). Previous studies on diboride

compounds gave a good working description of their chemical, physical and

thermodynamic properties in regard to their response to high temperature oxidizing

environments. Further tests showed that the addition of SiC to the ZrB2 and HfB2

composites would further improve their oxidation characteristics while not being

detrimental to the high temperature stability of these materials. This study also generated

mechanical property data for ZrB2 and HfB2 over a range of temperatures from 1200C to

inert atmospheres. Limited arc plasma tests showed favorable thermal stress and

oxidation resistance.

The Air Force program was divided into three phases: (I) composition and

microstructure screening, (II) extensive properties testing, and (III) simulated application

evaluations and verification of properties in a scaled-up fabrication. In the first phase,

oxidation, mechanical, and thermal screening tests were conducted for a variety of

compositions and microstructures in order to determine which materials should be further

studied. The second phase tested the thermal, physical, electrical and optical properties

for variations of ZrB2 and HfB2 with additions of SiC and SiC/C. The thermal

conductivity was measured from 100C to 1000C using the cut-bar method. All data

obtained showed a small negative temperature dependence that varied with the

composition and porosity of the material. The thermal diffusivity was measured from

1000C to 2000C using the flash-laser technique. Again all of the data showed slightly

negative temperature dependence. The thermal conductivity was also calculated over the

same temperature range from the diffusivity, specific heat, and density data. At 1000C

the conductivity calculated agreed with the data taken for the temperature range from

100C to 1000C and the small negative temperature dependence continued up through

2000C. In the third phase, material properties were found using scaled-up models.

Many application-oriented evaluations took place during this phase, including a test for a

leading edge configuration that was subjected to a simulated hypersonic flight heating.

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Purpose

Using the knowledge and experience gained in the development of supersonic

aircraft, the first hypersonic vehicle concepts were slender body vehicles with sharp

leading edges that would produce weak shock waves in order to minimize wave drag.

However, at hypersonic speeds aerothermodynamic heating is severe and sharp leading

edges made of traditional materials will naturally blunt through ablation to a larger radius

(Figures 3 and 4). Therefore, ultra-high temperature thermal protection system materials

would be needed in order to have non-ablating sharp leading edges.

NASA Ames Research Center has been developing UHTC’s for this explicit

purpose. UHTC’s are being investigated for use as a reusable thermal protection system

for flights beyond the capabilities that current state-of-the-art TPS materials being used

on the Space Shuttle Orbiter will allow. UHTC’s have exceptionally stable

configurations at temperatures of 1700C to 2800C with the presence of high velocity.

They also resist thermal shock and fatigue failure for operation over a multi-mission

lifecycle.

In order to minimize the mass of thermal protection systems, temperature profiles

are used to determine the selection and location of TPS materials on hypersonic vehicles.

The mass of the TPS system can be reduced by switching to a lighter material that has a

lower performance as the temperature decreases. The maximum non-ablating use

temperature of the next TPS material determines the location where the material change

should be made. A rigid surface TPS insulation that has been designed for high

temperature applications is the Alumina Enhanced Thermal Barrier (AETB) insulation.

AETB is shape stable up to temperatures of about 1426.7C. Above this temperature the

insulation begins to deform and soften under aerodynamic shear.

The transition from UHTC’s to AETB should be near where the surface

temperature is 1426.7C. The surface temperature will continue to decrease with distance

from the leading edge and therefore it may be advantageous to transition to yet another

material in order to further reduce mass and life-cycle costs. For most thermal protection

systems on sharp bodies the leading edge component will be a UHTC that will be

followed by a rigid surface insulation material.

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Testing

Arc jet testing of different variations of materials was conducted at the Aero

Heating Facility at Ames Research Center and was completed in 1997 (Figures 5 and 6).

The goal of these tests was to produce a material that had a thin, dense, and adherent

oxidation layer. For this to be accomplished, the tested samples were analyzed before

and after the arc jet testing to investigate the material performance in an aeroheating

environment (Selvaduray, 1997). The samples used for this series of testing were 0.75

inch in diameter and 0.25 inch thick. Some specimens also had a hole bored into them

that was 1/32 of an inch from the front surface. Fiber optic sensors were used to measure

the temperature in the hole and other sensors were used to measure the temperature of the

front surface.

On average, the oxide layer measured about 400 mm thick and has as high as a

300C temperature drop across them. The measurements on the back face showed a

temperature drop from 500 to 700C across the sample. The samples were analyzed using

petrography, X-Ray Diffraction, SEM, Energy Dispersive X-Ray, and other techniques as

well. The petrography techniques were developed in order to analyze the microstructure

of the samples before and after arc jet testing. An analysis of the microstructure showed

that during the testing process agglomerations formed between fine-grained particles.

The agglomerations caused highly porous regions that began failure and caused ablation

performance to be poor. It was determined at this point that further work needed to be

done on improving the manufacturing of these materials.

XRD analysis was able to provide useful information on the standard diffraction

patterns of the material and the back-surface stress analysis of post-test samples.

Diffraction patterns were taken of each sample prior to being tested to check for

impurities and unexpected phase transformations in the material that may have been

produced in the processing stage. Diffraction patterns were also taken on selected post-

test samples in order to analyze the change in the surface chemistry of the sample due to

the oxidation layer. In order to determine the back-surface stress on the samples a

different X-Ray tube was installed in the XRD. A powder sample was analyzed to

determine the baseline stress. After this the completed pre- and post-test samples were

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placed in the XRD unit to measure the back-surface stress due to the arc jet testing. The

back face of the samples was unexpectedly found to be in compression.

Flexural testing and analysis of UHTC materials was conducted using four-point

bending. The relevant strengths of different combinations of ZrB2, ZrC, HfB2/SiC, and

SiC were analyzed. The pre-test analysis for each sample consisted of the bulk density

and weight loss/gain measurements with sonic modulus testing. The sonic modulus is

used as a non-destructive technique to estimate Young’s Modulus by measuring the

natural frequency of the material and converting that into a modulus by taking into

account the dimensions and material properties of the sample. The flexural testing was

done using a computer controlled Instron 1122 testing machine. The post-test analysis

also involved the use of petrography on the samples after fracture.

By comparing the average strength, average density, and SiC agglomeration

formations it was concluded that the addition of SiC increased the average strength and

density of the specimen, but the porous agglomerations of SiC limited the strength

possible. For these tests the highest strength recorded was 108 ksi, but a properly

processed ZrB2/SiC can exhibit strengths of 140 ksi. However, these test specimens were

machined out of scrap sections and are presumed to be a worst-case scenario.

A graphite model holder was also designed for the following series of arc jet

testing. The purpose of these experiments is to measure the catalyticity of the UHTC

materials. The sample size was increased to 3 inches in diameter and the holes will be

bored within 1/16 of an inch from the front-face. A total of five sensors were implanted

as opposed to only one sensor in previous tests in order to allow in-depth and back-face

temperatures to be measured at the same time.

The focus for the TPS flight experiments are to demonstrate non-ablating sharp

leading edges, a molded structural TPS, and that Atmospheric-Assisted Ascent (AAA)

Technologies will enable a next generation reusable launch vehicle. The main objectives

for these tests are to show that sharp leading edges will minimize aerodynamic drag, life-

cycle cost, and mass of the structure.

The first flight test, the SHARP B-01 flight, took place in May of 1997. The main

objective for this flight was to show non-ablating performance of a 3-D UHTC nose tip

by performing a ballistic entry through the aerothermal performance constraint. The

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ground-based sensors were correlated with flight sensors to characterize the material

performance. The SHARP B-1 showed the performance of 1 mm radius leading edge at

Mach numbers greater than 22 and altitudes above 43 km. This flight test was a success,

however no data is available at this time.

The second flight test, SHARP-B02, was performed on September 28, 2000 as a

follow-on to the flight test performed in 1997 (Figure 7). Four UHTC leading edge

strakes were flown on a modified USAF Mk12 reentry vehicle. One pair of strakes was

retracted just before the temperatures were high enough to begin ablating. The other pair

was retracted shortly after ablation began. The SHARP B-2 reached its target point of

Kwajalein lagoon and was recovered on October 1, 2000. Specific results from this test

are still being analyzed and are not available at this time.

Maturity of Technology

Lincoln outlines five factors that he believes to be essential for successful

transition of a technology from the laboratory to full-scale development. The first factor

is stabilized material and/or material processes. This is an especially important factor

when dealing with UHTC’s because of the porous agglomerations observed in SiC during

the coupon tests. These agglomerations were amplified by poor processing procedures.

The conclusion from the flexural and arc jet testing was that further work needed to be

conducted on the processing of these materials in order to improve the uniformity of the

specimens (Selvaduray, 1997).

The second factor Lincoln stresses is the producibility of the material. The

material must be able to be reproduced in large, and if needed, detailed quantities. A

major concern is also the inspectability of the material through the manufacturing

process. UHTC’s have been developed for a full-scale model however; no information is

available on the specific manufacturing processes of these materials. One strict criterion

in the inspecting process will be the uniformity of the material in order to maintain the

highest possible strength and performance from the material.

The characterized mechanical properties are the third point that Lincoln considers

when evaluating technology transition. Testing is currently being done in order to fully

determine the mechanical properties of UHTC’s at the conditions to which the material

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would be exposed to while in service. The properties of most interest include the strength

of the material, its thermal conductivity, and the overall mass of the material in order to

allow for the lightest weight possible in the structure.

Another factor Lincoln looks at is the predictability of structural performance.

During the ground arc jet testing there were unexpected results of compression on the

back-face of the sample (Selvaduray, 1997). Therefore, the predictability of these

materials is not entirely characterized at this point and further testing is necessary in order

to do so.

The final point Lincoln makes is supportability. No information could be found

on the repair capability of UHTC’s, however for composites the process is extremely

detailed. One of the main goals associated with the application of UHTC’s is for the

ability to reuse the material without a large refurbishment effort. The repairs associated

with UHTC’s would only need to be at a conceptual level in order to be able to

successfully transition the material.

Overall this technology is not yet ready to be implemented at the full-scale level.

More testing needs to be done in order to determine the material properties and expected

results while in use. Also, specific instructions on manufacturing the material need to be

developed in order to ensure the desired performance for the material. Several

modifications also need to be made in order to increase safety factors.

Comparison

Currently in Germany work has been conducted by Daimler-Benz Aerospace

(Dasa) to develop a low cost ceramic matrix composite that can be applied to a wide

range of thermal protection systems (Trabandt, U., et.al. 1999). Dasa has developed a

CMC manufacturing process which allows the fabrication of complex and large integral

hot structures for applications such as nose caps, panels, and leading edges. Two TPS

concepts have been developed by Dasa, the first of which is the shingle concept. Here

small plane panels of C/SiC are mounted on a load carrying cold structure. The integral

hot structure concept allows for the possibility to fabricate conical and curved load-

carrying bodies to be applied for the thermal protection.

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One of the main objectives of the German program “Hot Structures” is to develop

new fabrication and joining techniques especially for hot structural CMC parts. This

program is based on a preliminary design of a small lifting body where the re-entry

trajectory can be controlled by fin rudders.

The Japanese are currently developing the HOPE-X, an experimental spaceplane

that will require a thermal protection system (Imuta and Gotoh, 1999). The HOPE-X is a

reentry-winged vehicle launched by the H-Ha rocket booster. Upon return to earth it will

land on a conventional runway.

During reentry, the temperature on the HOPE-X will reach 1700C. The Japanese

will be protecting the HOPE-X by using hot structures made of lightweight and heat

resistant materials. Development of these materials and a TPS are the two most

important technologies needed in developing the HOPE-X.

The concept of using new reusable lightweight materials is not unique to the

United States. Both Germany and Japan are actively researching this idea, to name only

two. Ultra-High Temperature Ceramics are quickly becoming the main focus for the

future of aerospace vehicles and thermal protection systems.

Conclusion

The development of Ultra-High Temperature Ceramics has had a revolutionary

impact on the design of thermal protection systems and reentry vehicles. Once these

materials can be implemented, significant increases in performance, life-cycle costs, and

a decrease in the weight of the TPS will be seen. This will make missions much more

economical by saving on fuel and repair costs because of thermal heating.

UHTC’s have been in the developmental process for over thirty years when the

Air Force began looking at ceramic matrix composites. These materials were seen to

obtain unique thermal qualities that were recognized for use in space applications.

Further testing of these materials has continued at NASA Ames Research Center in their

ground arc jet testing facilities. As recently as September of this year, there have been

flight tests in order to further understand the behavior of these composites.

There is still further testing to be completed before the material is ready to be

implemented in a full-scale design, but much has been achieved in the past thirty years.

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Other nations have also been designing their own version of UHTC’s for use on thermal

protection systems. Japan, for example, hopes to have the HOPE-X launched by early in

the 21st century with these “hot structures” in place as the TPS. The potential for

UHTC’s is overwhelming in comparison to what is currently in use and will most likely

have a radical impact on the future of the aerospace industry.

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References

Chen, Y-K, et al. “Integrated analysis tool for ultra-high temperature ceramic slender-

body reentry vehicles” AIAA, Aerospace Sciences Meeting and Exhibit (1999 ):

9 Sept 2000.

Clougherty, Edward V., Wilkes, Kenneth E., and Tye, Ronald P. Research and

Development of Refractory Oxidation-Resistant Diborides Part II, Vol. V

November, 1969. pp. 1-10.

Imuta, M., and Gotoh J. “Development of High Temperature Materials Including CMCs

for Space Application” Key Engineering Materials Vol. 164-165 (September,

1998) pp. 439-44.

Lincoln, John W. Structural Technology Transition to New Aircraft 14th Symposium of

the International Committee on Aeronautical Fatigue, Ottawa, Ontario, Canada,

pp. 1-10, 1987.

“SHARP-B2” 18 October, 2000 http://asm.arc.nasa.gov/projects/sharpb2/sharpb2.shtml

21 October, 2000.

“Sharp Body Aero-Space Vehicles, Using Ultra-High Temperature Ceramics” 2 pars. 9

Sept. 2000 http://ctoserver.arc.nasa.gov:80/techbreak/ceramics.html.

“SHARP: Slender Hypervelocity Aerothermodynamic Research Probes” 22 July, 2000

http://asm.arc.nasa.gov/projects/sharp/index.shtml 21 October, 2000.

“SHARP: Ultra-High Temperature Ceramics” 22 July, 2000

http://asm.arc.nasa.gov:80/projects/sharp/pl14.shtml 21 October, 2000.

Selvaduray, Guna, et al. “Development of Processing Techniques for Advanced Thermal

Protection Materials” 1997 pp. 11-13.

Trabandt, U., Wulz, H.G., and Schmid, T. “CMC for Hot Structures and Control Surfaces

of Future Launchers” Key Engineering Materials Vol. 164-165 (September, 1998)

pp. 445-449.

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Figures

Figure 1: Sharp bodies will allow for significant improvements in aerodynamic

performance for space travel.

Figure 2: UHTC’s will be used on the nose tips of reentry vehicles.

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Figure 3: In 1950's aerodynamics a blunt nose was used on aerospace vehicles in order to

create a bow shock to transfer kinetic energy to the boundary layer instead of the body.

Figure 4: Sharp leading edges reduce the aerodynamic drag and using 1990’s technology

the material has the capabilities to withstand substantial heating.

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Figure 5: Detailed view of a SHARP nose tip ready for testing in the arc jet heating

facilities at NASA Ames Research Center.

Figure 6: Arc jet testing of UHTC’s at NASA Ames Research Center

Figure 7: Flight testing on SHARP B-2 was a follow-on to SHARP B-1 and was

conducted in September of 2000.

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